A gas turbine shaft supports a series of disks. Each disk circumference supports a circular array of radially oriented aerodynamic blades. Closely surrounding these blades is a refractory shroud that encloses the flow of hot combustion gasses passing through the engine at temperatures of over 1400° C. The shroud is assembled from a series of adjacent rings supporting flow path components that are typically made of one or more refractory materials such as ceramics. Shroud rings that surround turbine blades are normally formed of a series of arcuate segments. Each segment is attached to a surrounding framework such as a metal ring called a blade ring that is, in turn, attached to the engine case. Close tolerances must be maintained in the gap between the turbine blade tips and the inner surfaces of the shroud ring segments to ensure engine efficiency. However, the shroud ring segments, blade ring, blades, disks, and their mountings are subject to differential thermal expansion during variations in engine operation, including engine restarts. This requires a larger gap and a corresponding efficiency reduction during some stages of engine operation.
Differences among coefficients of linear thermal expansion in flow path components and their support structures dictate the magnitude and variability of blade tip clearances. In prior designs, flow path components such as shroud ring segments are attached directly to support structures such as blade rings. Thus, when the support structures expand, the flow path components are pulled with them. This creates a large blade clearance requirement, partly because of the time delay between heating of flow path components and their more-insulated support structures.